The present invention relates generally to aerofoils for an axial flow turbo machine and in particular to improvements to aerofoils for axial flow compressors and turbines of gas turbine engines.
Axial flow turbo machines typically comprise a number of alternate stator and rotor rows in flow series. Both the rotor and stator rows comprise annular arrays of individual aerofoils. In the case of the stator rows the aerofoils comprise stator vanes and in the case of the rotor rows the aerofoils comprise blades mounted upon a rotor which rotates about a central axis. Typically in turbomachines the rotor and stator rows are arranged in pairs to form stages. For compressor stages the arrangement for each stage is typically rotor followed by stator, whilst for a turbine stage it is the opposite, namely stator followed by rotor. The individual stages, and aerofoils thereof, in use have an incremental effect on the flow of fluid through the stage giving rise to an overall resultant combined effect on the fluid flowing through the turbomachine. For a compressor the individual stages each incrementally increase the pressure of the flow through the stage. For a turbine the pressure decreases as energy is extracted from the flow through the stages to rotate and drive the turbine rotors.
In order to reduce the cost and weight of turbomachines it is desirable to reduce the number upstages and/or number of aerofoils in the rows of the stage, within a multi-stage axial flow turbomachine. In particular in gas turbine aeroengines, it is desirable to reduce the number upstages in the turbines and compressors. This requires the stage loading (i.e. effect each stage has on the flow therethrough) and thus the aerodynamic loading on the individual stages and aerofoils to be increased in order to maintain the same overall effect on the fluid flow through the turbomachine. Unfortunately as the aerodynamic loading increases the flow over the aerofoil surface tends to separate causing aerodynamic losses. This limits the stage loading that can be efficiently achieved.
In highly loaded turbine blades which operate at low Reynolds numbers, laminar boundary layer separation of the flow over the downstream rear portion of the suction surface cannot be avoided, and the blade is designed so that the separation and transition to turbulent boundary layer flow occurs before the trailing edge of the blade. Such high lift turbine aerofoil designs, the separation problems associated with them and a proposed means of addressing some of these problems are described in our UK patent application number GB9920564.3.
In highly loaded compressors, which often operate at high Reynolds numbers, fully turbulent boundary layer flows are present over the surfaces, and the blade is designed such that this turbulent layer does not separate from the aerofoil surface. If separation does occur then at the trailing edge there will be an open separation, in which the boundary layer does not reattach to the surface, resulting in high losses, increased flow deviation, reduced turning in the blade row and loss of pressure rise.
It is therefore desirable to provide an aerofoil in which the aerodynamic loading can be improved without significantly affecting the aerodynamic efficiency due to boundary layer separation and/or which offers improvements generally.
According to the present invention there is provided an exial flow turbo machine, the aerofoil having a span, a leading edge, a trailing edge and a cambered sectional profile comprising a pressure surface and a suction surface extending between the leading edge and trailing edge; characterised in that at least one aerofoil cross bleed passage is defined in the aerofoil, the passage extends from the pressure surface through the aerofoil to the suction surface.
Preferably the aerofoil is adapted in use to be highly loaded. The aerofoil may have a high lift profile.
Preferably an end of the at least one passage adjacent the suction surface is disposed generally at a location on the suction surface at which, in use, boundary layer separation from the suction surface would normally occur.
Preferably the at least one passage is arranged to provide, in use, a bleed from the pressure surface to the suction surface.
The at least one passage may be angled towards the trailing edge of the aerofoil. Preferably a portion of the passage adjacent to the suction surface is at a shallow angle relative to the suction surface. Furthermore the portion of the passage adjacent to the suction surface may be at an angle of less than 20xc2x0 to the suction surface.
Preferably the at least one passage comprises a plurality of passages disposed along the span of the aerofoil. The plurality of passages may be disposed in a row substantially parallel to the aerofoil span. Furthermore the plurality of passages may be disposed in at least two rows substantially parallel the aerofoil span. The passages of a first row of the at least two rows may also be staggered relative to the passages of a second row of the at least two rows.
The at least one passage may be curved as the passage extends from the pressure surface through the aerofoil to the suction surface.
The cross sectional area of the passage may vary as the passage extends from the pressure surface through the aerofoil to the suction surface. Preferably there is a portion of the passage adjacent to the suction surface, the cross sectional area of this portion of the passage decreases towards an end of the passage adjacent to the suction surface. Alternatively there is a portion of the passage adjacent to the suction surface, the cross sectional area of this portion of the passage increases towards an end of the passage adjacent to the suction surface.
Preferably the at least one passage comprises a slot extending along at least part of the aerofoil span and extending through the aerofoil from the leading to the trailing edge.
The at least one passage may comprise a first portion adjacent to the suction surface and a second portion adjacent to the pressure surface, the first portion extending through the aerofoil at an angle to the second portion. The at least one passage may comprise a plurality of passages disposed along the span of the aerofoil and the second portion of the passages comprises a slot common to at least two of the passages and extending along at least part of the aerofoil span.
Preferably the aerofoil comprises part of a blade for a turbo machine. Alternatively the aerofoil may comprise part of a vane for a turbo machine.
The aerofoil may comprise a compressor aerofoil. The aerofoil profile may have a thickness between the pressure and suction surfaces, which increases from the leading edge to a maximum thickness at a position along a chord of the aerofoil closer to the trailing edge than to the leading edge. The maximum thickness of the aerofoil is preferably at a position from the leading edge substantially two thirds of the way along chord. An end of the at least one passage adjacent the suction surface may be disposed generally downstream of the position of maximum thickness of the aerofoil. Preferably an end of the at least one passage adjacent the suction surface is disposed generally downstream of the position of maximum curvature of the aerofoil.
The aerofoil may comprise a turbine aerofoil. An end of the at least one passage adjacent to the pressure surface may be disposed generally in a region of the pressure surface extending from the leading edge where, in use, boundary layer separation from the pressure surface would normally occur.
Preferably the at least one passage has a generally circular cross section. Alternatively the at least one passage may have a generally elliptical cross section.
The aerofoil may comprise part of a gas turbine engine.